NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Reynolds number: 200,000 Max Cl/Cd: 41.26 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rcsc2-il-200000-n5.txt Download as CSV file: xf-rcsc2-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-SC2 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.6413 0.09110 0.08689 -0.0290 1.0000 0.0261
-10.750 -0.7393 0.06221 0.05776 -0.0504 1.0000 0.0249
-10.500 -0.7893 0.05463 0.04989 -0.0505 1.0000 0.0247
-10.250 -0.8251 0.05051 0.04553 -0.0455 1.0000 0.0246
-10.000 -0.8489 0.04671 0.04143 -0.0407 1.0000 0.0245
-9.750 -0.8633 0.04344 0.03784 -0.0363 1.0000 0.0246
-9.500 -0.8706 0.04058 0.03468 -0.0322 1.0000 0.0247
-9.250 -0.8725 0.03804 0.03184 -0.0285 1.0000 0.0248
-9.000 -0.8700 0.03575 0.02924 -0.0252 1.0000 0.0250
-8.750 -0.8641 0.03368 0.02688 -0.0222 1.0000 0.0252
-8.500 -0.8553 0.03180 0.02472 -0.0195 1.0000 0.0254
-8.250 -0.8441 0.03010 0.02276 -0.0170 1.0000 0.0257
-8.000 -0.8308 0.02864 0.02104 -0.0148 1.0000 0.0262
-7.750 -0.8160 0.02728 0.01941 -0.0127 1.0000 0.0267
-7.500 -0.7997 0.02606 0.01793 -0.0108 1.0000 0.0272
-7.250 -0.7822 0.02496 0.01661 -0.0090 1.0000 0.0275
-7.000 -0.7640 0.02377 0.01536 -0.0075 1.0000 0.0278
-6.750 -0.7451 0.02283 0.01435 -0.0061 1.0000 0.0282
-6.500 -0.7257 0.02200 0.01347 -0.0047 1.0000 0.0286
-6.250 -0.7059 0.02123 0.01265 -0.0033 1.0000 0.0291
-6.000 -0.6857 0.02051 0.01188 -0.0020 1.0000 0.0297
-5.750 -0.6567 0.01976 0.01105 -0.0025 0.9983 0.0304
-5.500 -0.6275 0.01905 0.01028 -0.0031 0.9965 0.0313
-5.250 -0.5982 0.01843 0.00959 -0.0036 0.9949 0.0322
-5.000 -0.5704 0.01781 0.00898 -0.0039 0.9932 0.0334
-4.750 -0.5425 0.01735 0.00853 -0.0042 0.9911 0.0351
-4.500 -0.5134 0.01690 0.00806 -0.0046 0.9890 0.0370
-4.250 -0.4835 0.01643 0.00757 -0.0053 0.9869 0.0389
-4.000 -0.4526 0.01604 0.00719 -0.0061 0.9851 0.0415
-3.750 -0.4233 0.01570 0.00683 -0.0066 0.9830 0.0451
-3.500 -0.3946 0.01533 0.00650 -0.0069 0.9797 0.0510
-3.250 -0.3633 0.01496 0.00619 -0.0078 0.9768 0.0607
-3.000 -0.3309 0.01460 0.00592 -0.0090 0.9744 0.0762
-2.750 -0.2982 0.01425 0.00571 -0.0102 0.9726 0.1009
-2.500 -0.2726 0.01389 0.00553 -0.0100 0.9685 0.1354
-2.250 -0.2434 0.01346 0.00536 -0.0105 0.9651 0.1887
-2.000 -0.2130 0.01282 0.00519 -0.0115 0.9622 0.2916
-1.750 -0.1857 0.01175 0.00506 -0.0121 0.9600 0.4974
-1.500 -0.1647 0.01109 0.00511 -0.0105 0.9555 0.6539
-1.250 -0.1384 0.01086 0.00522 -0.0096 0.9516 0.7429
-1.000 -0.1073 0.01081 0.00534 -0.0098 0.9488 0.8013
-0.750 -0.0727 0.01087 0.00553 -0.0104 0.9467 0.8537
-0.500 -0.0392 0.01099 0.00572 -0.0109 0.9435 0.8934
-0.250 -0.0036 0.01109 0.00584 -0.0119 0.9377 0.9174
0.000 0.0476 0.01106 0.00580 -0.0159 0.9326 0.9348
0.250 0.0886 0.01095 0.00569 -0.0179 0.9214 0.9469
0.500 0.1294 0.01079 0.00553 -0.0199 0.9103 0.9546
0.750 0.1686 0.01061 0.00534 -0.0218 0.8993 0.9598
1.000 0.2092 0.01043 0.00516 -0.0241 0.8857 0.9627
1.250 0.2452 0.01025 0.00497 -0.0253 0.8635 0.9667
1.500 0.2827 0.00997 0.00462 -0.0266 0.8231 0.9699
1.750 0.3191 0.00984 0.00423 -0.0275 0.7475 0.9724
2.000 0.3516 0.01013 0.00404 -0.0278 0.6426 0.9749
2.250 0.3784 0.01078 0.00410 -0.0276 0.5144 0.9793
2.500 0.4029 0.01159 0.00429 -0.0272 0.3851 0.9847
2.750 0.4333 0.01221 0.00450 -0.0282 0.2973 0.9881
3.000 0.4635 0.01275 0.00471 -0.0291 0.2303 0.9920
3.250 0.4939 0.01330 0.00491 -0.0301 0.1705 0.9954
3.500 0.5260 0.01374 0.00513 -0.0314 0.1294 0.9986
3.750 0.5518 0.01410 0.00536 -0.0313 0.1050 1.0000
4.000 0.5719 0.01443 0.00559 -0.0299 0.0883 1.0000
4.250 0.5918 0.01476 0.00585 -0.0285 0.0753 1.0000
4.500 0.6115 0.01511 0.00613 -0.0270 0.0643 1.0000
4.750 0.6311 0.01548 0.00645 -0.0255 0.0552 1.0000
5.000 0.6506 0.01586 0.00683 -0.0239 0.0486 1.0000
5.250 0.6697 0.01629 0.00722 -0.0222 0.0440 1.0000
5.500 0.6890 0.01670 0.00765 -0.0205 0.0406 1.0000
5.750 0.7076 0.01719 0.00810 -0.0188 0.0382 1.0000
6.000 0.7265 0.01765 0.00861 -0.0171 0.0363 1.0000
6.250 0.7455 0.01812 0.00910 -0.0154 0.0345 1.0000
6.500 0.7640 0.01864 0.00961 -0.0137 0.0331 1.0000
6.750 0.7822 0.01924 0.01022 -0.0120 0.0319 1.0000
7.000 0.8008 0.01984 0.01088 -0.0102 0.0309 1.0000
7.250 0.8192 0.02048 0.01157 -0.0085 0.0302 1.0000
7.500 0.8375 0.02116 0.01229 -0.0068 0.0295 1.0000
7.750 0.8560 0.02187 0.01304 -0.0052 0.0289 1.0000
8.000 0.8744 0.02263 0.01383 -0.0036 0.0284 1.0000
8.250 0.8927 0.02348 0.01471 -0.0021 0.0279 1.0000
8.500 0.9108 0.02451 0.01576 -0.0006 0.0275 1.0000
8.750 0.9296 0.02551 0.01687 0.0008 0.0272 1.0000
9.000 0.9484 0.02657 0.01807 0.0022 0.0269 1.0000
9.250 0.9666 0.02773 0.01938 0.0037 0.0267 1.0000
9.500 0.9840 0.02900 0.02082 0.0052 0.0264 1.0000
9.750 1.0004 0.03038 0.02240 0.0067 0.0261 1.0000
10.000 1.0154 0.03189 0.02412 0.0085 0.0259 1.0000
10.250 1.0286 0.03354 0.02599 0.0103 0.0257 1.0000
10.500 1.0396 0.03532 0.02802 0.0124 0.0256 1.0000
10.750 1.0483 0.03720 0.03016 0.0146 0.0254 1.0000
11.000 1.0542 0.03917 0.03238 0.0171 0.0251 1.0000
11.250 1.0572 0.04122 0.03467 0.0197 0.0249 1.0000
11.500 1.0553 0.04327 0.03695 0.0229 0.0247 1.0000
11.750 1.0501 0.04550 0.03939 0.0261 0.0246 1.0000
12.000 1.0419 0.04802 0.04214 0.0290 0.0244 1.0000
12.250 1.0296 0.05103 0.04538 0.0313 0.0243 1.0000
12.500 1.0117 0.05486 0.04947 0.0328 0.0243 1.0000
12.750 0.9847 0.06027 0.05517 0.0325 0.0243 1.0000
13.000 0.9372 0.07030 0.06560 0.0270 0.0245 1.0000
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Polar data table (+)
Polar graphs
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